This invention relates generally to a cooling system for a gas turbine engine component, and, more particularly, to a cooling system for an electronic engine control.
The sophistication of gas turbine engines has developed to the utilization of electronic engine controls to supplement, and in some cases replace, hydromechanical engine controls for providing improved safety and higher efficiency of operation. However, an electronic control module is more sensitive to temperature than is a hydromechanical control and therefore requires a greater degree of cooling to maintain its reliability in normal operation and to prolong its useful service life.
Conventional cooling systems are relatively complex and may adversely affect the overall efficiency of the engine. Various heat sink sources have been utilized and include singly or in combination freestream or outside airflow, fan or compressor bleed air and even engine fuel. An example of a conventional manner of keeping the module cool on a gas turbine engine, such as a turboshaft engine, is to install the module in an airframe nacelle or on an engine frame, that is, the external, metallic structure of the engine, where the temperature is relatively low. On a turbofan engine, the module can be installed in the annular airspace within the nacelle between the fan casing and the nacelle where, because this portion of the nacelle is spaced away from the engine core, the temperature is lower than it is immediately adjacent the core.
However, the temperature in the airframe nacelle of a gas turbine engine is still too warm to enable best, long-life operation unless supplemental cooling is utilized. Such supplemental cooling can be obtained by blowing air across the module, and is most advantageous if the coolest air available is utilized to obtain the greatest amount of cooling.
The source of air commonly used for cooling engine components such as an electronic control module is air bled from the initial stages of the compressor of the engine, or, in a turbofan engine, fan air from behind the fan. The air from each of these sources has pressure increased by the compressor or fan, and is thus warmer and consequently a less desirable source of cooling air than is the unpressurized, unheated outside freestream air surrounding the engine nacelle. Furthermore, if ram air, that is, the freestream air which is forced into an aircraft engine as the aircraft moves through the air, is utilized for cooling, there is no flow of air when the aircraft is stationary on the ground.
Another conventional manner of cooling an engine electronic control is disclosed in U.S. Pat. No. 4,351,150--W. M. Schulze, assigned to the present assignee. The auxiliary cooling air system disclosed therein represents an improvement over prior art systems. However, the Schulze system is relatively complex and includes additional air plumbing and a jet pump which also uses compressor bleed air for moving outside air over the electronic control. Of course, it will be appreciated that the use of bleed air in any cooling system reduces the overall gas turbine engine efficiency.
Yet other conventional systems for cooling electronic components, in general, include various heat dissipating fin structures, some of which extend into a flow channel through which cooling air is driven by an auxiliary fan. However, the use of fin structures disposed directly in the flowpath inside a gas turbine engine is not known, inasmuch as any obstructions in the flowpath might adversely affect the desired aerodynamic flow patterns in the inlet to the fan or compressor, for example.
Another important function in the efficient operation of a gas turbine engine involves the determination of engine inlet air temperature and pressure. The thrust or shaft horsepower developed by a gas turbine engine and the engine control settings are dependent in part upon the temperature and pressure of the air entering the engine. Consequently, this inlet air must be measured to adjust the fuel flow into the engine for obtaining the desired output power.
Inlet air temperature and pressure sensors are commonly located at positions on the engine nacelle upstream of the engine compressor, and upstream of the fan in the case of a turbofan engine, such that the sensors are directly exposed to the engine inlet airstream. However, this placement can result in inaccurate readings or even loss of temperature and pressure measuring capability. For example, the sensors may accumulate an ice coating at some atmospheric conditions, or they may experience foreign object damage from bird strikes or earth particles impinging upon the sensors.
An even more serious difficulty may arise if the sensors are located on the inner surface of the engine inlet cowl. If a sensor, or part of one, breaks off, as could occur, for example, during a bird strike, the loose piece would be ingested by the fan or compressor and could cause serious damage or even lead to engine failure. Therefore, anti-icing systems and foreign object protection devices are conventionally utilized for protecting these measurement devices.
Accordingly, it is an object of the present invention to provide a new and improved component cooling system for a gas turbine engine.
Another object of the present invention is to provide an electronic engine control cooling system which does not require an auxiliary air source or bleed air for providing cooling.
Another object of the present invention is to provide a cooling system which is relatively simple and which uses engine inlet air as a cooling fluid.
Another object of the present invention is to provide a cooling system which directly incorporates and protects temperature and pressure measurement sensors.